1. Field of the Invention
The present invention relates to a cooled turbine blade for a gas turbine and more particularly to a cooled turbine blade provided therein with cooling flow passages and elements related thereto.
2. Description of the Related Art
There are various types of gas turbines, and one of them is of a direct driving type which is driven by a combustion gas flow which drives a compressor. The higher the temperature of the main flow gas is, the higher heat efficiency the direct driving gas turbine has. Thus, attempts have been made to increase the temperature of the main gas flow. In general, however, the upper temperature of the main gas flow is limited by the heat resistance of the turbine blade. In order to improve the heat efficiency by increasing the temperature of the main gas flow, therefore, a turbine blade which withstands a higher temperature is required. The heat resistance of the turbine blades is enhanced by improving the material of the turbine blade and cooling the turbine blades from their inside to lower their surface temperatures. The turbine blade of this type is formed therein with cooling flow passages for conducting water or a cooling gas such as water vapor or air and uses various means for increasing cooling efficiency.
A conventional cooled type turbine blade of this type is shown in FIGS. 31 and 32. FIG. 31 is a transverse cross-sectional view of the turbine blade and FIG. 32 is a longitudinal cross-sectional view thereof. The turbine blade 1 has an aerofoil blade portion 2 in which a plurality of serially arranged cooling flow passages 4, 10, 12 and 13 are formed extending in the span direction. A cooling gas flows through the cooling flow passages 12, 10 and 13 via passages formed in a shank 3 to cool the wall portions 6 and 7 of the blade. A great number of nozzles 8 and 17 are formed in the wall portions 6 and 7, and part of the cooling gas flowing through the cooling flow passages 4 and 13 is jetted out of the nozzles 8 and 17. The jetted cooling gas flows in a film state along the surfaces of the suction side and the pressure side of the aerofoil blade portion 2 so as to interrupt heat transmitted from the environmental main gas flow to the surface of the aerofoil blade portion 2 and so as to cool the surface of the aerofoil blade portion 2. Thus, so-called film cooling is performed.
An impingement chamber or a leading edge chamber is formed in the leading edge portion of the blade. The cooling gas supplied to the cooling flow passage 4 is jetted from a great number of small holes and impinges on the inner surface of the leading edge wall 5, whereby performing so-called impingement cooling. The leading edge wall 5 is formed with a great number of nozzles to form a so-called shower head 9. The cooling gas in a leading edge chamber is jetted from the washer head 9 and the film cooling is performed.
In a trailing edge portion 15 of the blade is formed a trailing edge chamber into which the cooling gas flows from the cooling flow passage 13 through a nozzle 14. In the trailing edge of the blade is formed a slit-shaped trailing edge nozzle 16 from which the cooling gas in the trailing edge chamber is exhausted externally. A great number of the pin fins 11 are formed in the trailing edge chamber to improve the cooling efficiency of the trailing edge portion 15.
With the cooled turbine provided with such a turbine blade, the average surface temperature of the blade can be maintained at 850.degree. C. when the temperature of the main flow gas is within a range from 1,000.degree. C. to 1,300.degree. C. In this case, the amount of flow of the cooling gas is several percent of that of the main flow gas. Recently, however, a gas turbine has been developed that operates at a main flow gas temperature from 1,300.degree. C. to 1,500.degree. C. Further, development of a gas turbine which uses hydrogen as a fuel and operates at a temperature from 1,500.degree. C. to 2,000.degree. C. is now under consideration.
If such an improved gas turbine were manufactured on the basis of the design of the conventional gas turbine, the amount of cooling gas would have to be made extremely large in order to maintain an average surface temperature of 850.degree. C., and thus the heat efficiency of a gas turbine or the whole heat plant including the gas turbine would be extremely reduced. Therefore, it is difficult to actualize such an improved gas turbine.
A cooled turbine blade was proposed which can provide a higher cooling efficiency without increasing the flowing amount of the cooling gas, which is disclosed, for example, in the U.S. Pat. No. 5,165,852. The cooled turbine blade is provided in its aerofoil blade portion with a first cooling flow passage disposed at the pressure side and a second cooling flow passage disposed at the suction side. The cooling gas flows in the radial outward direction of the turbine rotor including the turbine blade, i.e., from the shank of the turbine blade toward the wing tip portion in the first cooling flow passage and in the radial inward direction, i.e., from wing tip portion toward the shank in the second cooling flow passage.
When the cooling gas flows in the radial direction of the turbine rotor, a Coriolis force produced by rotation of the rotor is exerted on the cooling gas. A secondary flow in the direction crossing the cooling flow passage occurs in the cooling gas flowing through the cooling flow passage, and a pair of longitudinal vortexes are produced in the cooling gas flowing through the cooling flow passage. The vortexes in the first cooling flow passage and the second cooling flow passage are directed opposite to each other. The cooling gas collides against the inner surface of the pressure side wall of the aerofoil blade portion in the first cooling flow passage and on the inner surface of the suction side wall of the aerofoil blade portion in the second cooling flow passage. In this way, both the pressure side and the suction side of the aerofoil blade portion are effectively cooled and thus a high cooling efficiency can be attained by a small amount of the cooling gas.
One of means for effectively cooling the wall of the cooling flow passage (i.e., the wall portion of the aerofoil blade portion) by the flow of the cooling gas in the cooling flow passage comprises a great number of projections formed in the inner surface of the wall of the cooling flow passage so that the cooling gas flows in a turbulent flow state in the vicinity of the inner surface of the wall.
An embodiment of a turbine blade having such projections is shown in FIGS. 33 and 34. The turbine blade 21 has a shank portion 21a and an aerofoil blade portion 21b. In the aerofoil blade portion 21b are formed a plurality of cooling flow passages 22 having their ends serially connected together by means of return portions 24. A cooling gas supplied from a cooling gas inlet 23 flows through the adjacently arranged cooling flow passages 22 via the return portions 24 and is finally discharged from a nozzle formed in a trailing edge portion 26 into the main gas.
On the walls of the cooling flow passages 22, for example, the inner surface of the suction side walls of the aerofoil blade portion are formed a plurality of turbulence promoting ribs 27 which extend perpendicularly to the flow direction of the cooling gas. The cooling gas is formed into a strong turbulence in the vicinity of the inner surface of the wall by means of the turbulence promoting ribs 27. The heat transmission through the inner surface is enhanced to improve the cooling effect. Partition walls 29 are intended to divide adjacent cooling flow passages 22, and pin fins 29 are formed in the trailing edge portion.
Jpn. Pat. KOKAI Publication No. 60-101202 discloses a blade having turbulence promoting ribs formed slantwise with respect to the flowing direction of the cooling gas. Such turbulence promoting ribs produce a strong turbulence to improve the cooling effect. The inclined turbulence promoting ribs can prevent foreign matters from being deposited on or at a specific portion. Jpn. Pat. KOKAI Publication No. 5-10101 discloses an arrangement of v-shaped slantwise turbulence promoting ribs.
With the gas turbine operating at 1,300.degree. C. to 1,500.degree. C. or at 1,500.degree. C. to 2,000.degree. C., it is necessary that the turbine blade be cooled at a higher cooling efficiency at a smaller flow rate of the cooling gas. Thus, it is required that not only the cooling efficiency of a plurality of cooling means be enhanced as described above but also a higher cooling efficiency be attained by the combination of such cooling means.
When the cooling flow passages extend in the span direction of the turbine blade, i.e., in the radial direction of the turbine rotor, they act as a kind of centrifugal fan due to a centrifugal force produced by a high rotational speed of the rotor. Upon increasing the cooling efficiency by a Coriolis force as described above, the cooling gas is supplied to the pressure side cooling flow passages in the radial outward direction and to the suction side passages in the radial inward direction. The cooling gas flow through the pressure side cooling flow passages is accelerated because the cooling gas flows in the radial outward direction. On the other hand, the cooling gas flow through the suction side passage is blocked because the cooling gas flows in the radial inward direction.
In the gas turbine blade as disclosed in the U.S. Pat. No. 5,165,852, the cooling flow passages at both the suction side and the pressure side, which are connected in series are the same in number so that the disadvantage of the flow of the suction side cooling gas is canceled out. In generally, the most downstream cooling flow passage communicates with the nozzle formed in the suction side or the pressure side of the aerofoil blade portion or the trailing edge portion. The cooling gas which has flowed through the most downstream coiling passage is discharged from the nozzle into the main flow gas to perform film cooling. In this structure, the most downstream passage constitutes the suction side cooling flow passage. Thus, the cooling gas in the suction side cooling flow passage is jetted out. Since the flow of the cooling gas in the suction side cooling flow passage is directed radially inward, the gas flow is also restricted. More specifically, the static pressure of the cooling gas in the suction side cooling flow passage is low. In this connection, the cooling gas must be supplied at a higher pressure in order to supply a predetermined amount of the cooling gas to the man flow gas through the nozzle. Increase of the driving power necessary for the cooling gas supply lowers the total heat efficiency of the gas turbine.
If the number of the cooling flow passages at the pressure side is made larger by one than those a the suction side, the most downstream cooling flow passage constitutes a pressure side passage. Thus, the increase of the driving power is prevented.
In general, the cross section of the aerofoil blade has a large camber, and thus the dimension of the suction side flow passages in the chord direction is larger than that of the pressure side flow passages in the chord direction. Provision of one more cooling flow passage at the pressure side than those at the suction side makes large the difference of the cross section between the cooling flow passages at the suction side and at the pressure side. In the cooling flow passages at the suction side, the gas velocity tends to decrease and to lower the cooling efficiency. Thus, it is not preferable that the cross section of the passages at the suction side be made larger than that at the pressure side.
It is possible to correct the imbalance of the cross sections by displacing the partition for dividing the interior of the aerofoil blade portion into the cooling flow passages at the suction side and at the pressure side. In this arrangement, however, the cooling flow passages has a flat cross section extending in the chord direction. Such flat cross section weakens longitudinal vortexes produced by a Coriolis force in the suction side cooling flow passages, leading to an unfavorable result that the cooling efficiency in the suction side cooling flow passages is much more reduced.
When the means for improving the cooling efficiency by a Coriolis force is used, there occurs the disadvantages in that the driving force required for conducting the cooling gas as described above and/or the cooling efficiency of the suction side cooling flow passages is lowered more. When, on the other hand, both means for improving the cooling efficiency by a Coriolis force and the turbulence promoting ribs are used, the pressure loss of the cooling gas produced by the turbulence promoting ribs is large and the above-mentioned disadvantages become much more remarkable. Further, turbulence produced by the turbulence promoting ribs interferes with longitudinal vortexes produced by the Coriolis force, whereby the multiplier effect of the turbulence promoting ribs and the Coriolis force is sometimes reduced.